Gas turbine engine shaft bearing configuration

ABSTRACT

A gas turbine engine includes a core housing that includes an inlet case and an intermediate case that respectively provide an inlet case flow path and an intermediate case flow path. A geared architecture is arranged within the inlet case. A shaft provides a rotational axis. A hub is operatively supported by the shaft. A rotor is connected to the hub and supports a compressor section. The compressor section is arranged axially between the inlet case flow path and the intermediate case flow path. A bearing is mounted to the hub and supports the shaft relative to one of the intermediate case and the inlet case.

CROSS REFERENCE TO RELATED APPLICATIONS

This application claims priority to provisional application No.61/860,337 filed Jul. 31, 2013, and this application is acontinuation-in-part of U.S. application Ser. No. 13/904,416 filed onMay 29, 2013, which is a continuation of U.S. application Ser. No.13/762,970 filed on Feb. 8, 2013, now U.S. Pat. No. 8,511,061 issuedAug. 20, 2013, which is a continuation of U.S. application Ser. No.13/362,170 filed on Jan. 31, 2012, now U.S. Pat. No. 8,402,741 issuedMar. 26, 2013.

BACKGROUND

Turbomachines, such as gas turbine engines, typically include a fansection, a turbine section, a compressor section, and a combustorsection. The fan section drives air along a core flow path into thecompressor section. The compressed air is mixed with fuel and combustedin the combustor section. The products of combustion are expanded in theturbine section.

A typical jet engine has two or three spools, or shafts, that transmittorque between the turbine and compressor sections of the engine. Eachof these spools is typically supported by two bearings. One bearing, forexample, a ball bearing, is arranged at a forward end of the spool andis configured to react to both axial and radial loads. Another bearing,for example, a roller bearing is arranged at the aft end of the spooland is configured to react only to radial loads. This bearingarrangement fully constrains the shaft except for rotation, and axialmovement of one free end is permitted to accommodate engine axialgrowth.

A core inlet typically controls flow of air into the core flow path. Theflow of air moves from the core inlet to a compressor section inlet. Therelative radial positions of the core inlet and the compressor sectioninlet may influence flow through the core and a profile of theturbomachine.

SUMMARY

In one exemplary embodiment, a gas turbine engine includes a corehousing that includes an inlet case and an intermediate case thatrespectively provide an inlet case flow path and an intermediate caseflow path. A geared architecture is arranged within the inlet case. Ashaft provides a rotational axis. A hub is operatively supported by theshaft. A rotor is connected to the hub and supports a compressorsection. The compressor section is arranged axially between the inletcase flow path and the intermediate case flow path. A bearing is mountedto the hub and supports the shaft relative to one of the intermediatecase and the inlet case.

In a further embodiment of the above, a core inlet includes the inletcase flow path and has a radially inner boundary that is spaced a firstradial distance from the rotational axis. A compressor section inlet hasa radially inner boundary that is spaced a second radial distance fromthe rotational axis. A ratio of the second radial distance to the firstradial distance is of about 0.65 to about 0.9.

In a further embodiment of the above, the radially inner boundary of thecore inlet is at a location of a core inlet stator.

In a further embodiment of the above, the radially inner boundary of thecompressor section inlet is at a location of a compressor rotor.

In a further embodiment of the above, the compressor rotor is a firststage rotor of a low-pressure compressor.

In a further embodiment of the above, the core inlet is an inlet to thecore housing.

In a further embodiment of the above, an inlet flow of the compressorsection is configured to be from about 30 lb/sec/ft² to about 37lb/sec/ft² when the engine is operating at a cruise speed.

In a further embodiment of the above, a turbine inlet temperature of ahigh-pressure turbine within the engine is configured to be from about2,000° F. to about 2,600° F. when the engine is operating at a cruisespeed.

In a further embodiment of the above, a tip speed of a blade array inthe compressor section during engine operation is configured to be fromabout 1,050 fps to about 1,350 fps.

In a further embodiment of the above, a fan section is driven by thegeared architecture that is driven by the shaft that rotates acompressor rotor within the compressor section.

In a further embodiment of the above, the geared architecture has a gearreduction ratio of greater than about 2.3.

In a further embodiment of the above, the inlet case includes a firstinlet case portion that defines the inlet case flow path. A bearingsupport portion is removably secured to the inlet case portion. Thebearing is mounted to the bearing support portion.

In a further embodiment of the above, the intermediate case includes anintermediate case portion that defines the intermediate case flow path.A bearing support portion is removably secured to the intermediate caseportion. The bearing is mounted to the bearing support portion.

In a further embodiment of the above, the bearing is a ball bearing.

In a further embodiment of the above, the bearing is a first bearing andfurther comprising a second bearing that supports the shaft relative tothe other of the intermediate case and the inlet case.

In a further embodiment of the above, the first and second bearings arearranged in separate sealed lubrication compartments.

In a further embodiment of the above, the geared architecture is coupledto the shaft. A fan is coupled to and rotationally driven by the gearedarchitecture.

In a further embodiment of the above, the shaft includes a main shaftand a flex shaft. The flex shaft is secured to the main shaft at a firstend and including a second end opposite the first end. The gearedarchitecture includes a sun gear supported on the second end.

In a further embodiment of the above, the shaft includes a hub that issecured to the main shaft. The compressor section includes a rotormounted to the hub.

In a further embodiment of the above, the geared architecture includes atorque frame that supports multiple circumferentially arranged stargears that intermesh with the sun gear. The torque frame is secured tothe inlet case.

In a further embodiment of the above, the rotor supports multiplecompressor stages. The bearing is axially aligned with and radiallyinward of one of the compressor stages.

In a further embodiment of the above, the compressor section includes avariable stator vane array.

In a further embodiment of the above, the geared architecture isarranged in the lubrication compartment.

In a further embodiment of the above, the core housing includes a coreinlet stator that has a stator root that is spaced a first radialdistance from the rotational axis. The compressor section includes acompressor blade having a blade root that is spaced a second radialdistance from the rotational axis. A ratio of the second radial distanceto the first radial distance is of about 0.65 to about 0.9.

In a further embodiment of the above, the stator root is radiallyaligned with a radially inner boundary of a core flow path through thegas turbine engine.

In a further embodiment of the above, the blade root is radially alignedwith a radially inner boundary of a core flow path through the gasturbine engine.

In a further embodiment of the above, the core inlet stator ispositioned within an inlet to a core section of the gas turbine engine.

BRIEF DESCRIPTION OF THE DRAWINGS

The disclosure can be further understood by reference to the followingdetailed description when considered in connection with the accompanyingdrawings wherein:

FIG. 1 schematically illustrates an embodiment of a gas turbine engine.

FIG. 2 is a cross-sectional view of a front architecture of the gasturbine engine embodiment shown in FIG. 1.

FIG. 3 shows a close-up view of a core inlet portion of the FIG. 1 gasturbine engine embodiment.

The embodiments, examples and alternatives of the preceding paragraphs,the claims, or the following description and drawings, including any oftheir various aspects or respective individual features, may be takenindependently or in any combination. Features described in connectionwith one embodiment are applicable to all embodiments, unless suchfeatures are incompatible.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flowpath B whilethe compressor section 24 drives air along a core flowpath C (as shownin FIG. 2) for compression and communication into the combustor section26 then expansion through the turbine section 28. Although depicted as atwo-spool turbofan gas turbine engine in the disclosed non-limitingembodiment, it should be understood that the concepts described hereinare not limited to use with two-spool turbofans as the teachings may beapplied to other types of turbine engines including three-spoolarchitectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a low pressure compressor 44 and a low pressureturbine 46. The inner shaft 40 is connected to the fan 42 through aspeed change mechanism, which in exemplary gas turbine engine 20 isillustrated as a geared architecture 48 to drive the fan 42 at a lowerspeed than the low speed spool 30. The high speed spool 32 includes anouter shaft 50 that interconnects a high pressure compressor 52 and highpressure turbine 54. A combustor 56 is arranged in exemplary gas turbine20 between the high pressure compressor 52 and the high pressure turbine54. A mid-turbine frame 57 of the engine static structure 36 is arrangedgenerally between the high pressure turbine 54 and the low pressureturbine 46. The mid-turbine frame 57 supports one or more bearingsystems 38 in the turbine section 28. The inner shaft 40 and the outershaft 50 are concentric and rotate via bearing systems 38 about theengine central longitudinal axis A, which is collinear with theirlongitudinal axes.

The core airflow C is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path. The turbines 46, 54 rotationally drive therespective low speed spool 30 and high speed spool 32 in response to theexpansion. It will be appreciated that each of the positions of the fansection 22, compressor section 24, combustor section 26, turbine section28, and fan drive gear system 48 may be varied. For example, gear system48 may be located aft of combustor section 26 or even aft of turbinesection 28, and fan section 22 may be positioned forward or aft of thelocation of gear system 48.

The engine 20 in one example a high-bypass geared aircraft engine. In afurther example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than ten (10). The examplespeed reduction device is a geared architecture 48 however other speedreducing devices such as fluid or electromechanical devices are alsowithin the contemplation of this disclosure. The example gearedarchitecture 48 is an epicyclic gear train, such as a star gear systemor other gear system, with a gear reduction ratio of greater than about2.3, or more specifically, a ratio of from about 2.2 to about 4.0. Inone disclosed embodiment, the engine 20 bypass ratio is greater thanabout ten (10:1), the fan diameter is significantly larger than that ofthe low pressure compressor 44, and the low pressure turbine 46 has apressure ratio that is greater than about 5:1. Low pressure turbine 46pressure ratio is pressure measured prior to inlet of low pressureturbine 46 as related to the pressure at the outlet of the low pressureturbine 46 prior to an exhaust nozzle. It should be understood, however,that the above parameters are only exemplary of one embodiment of ageared architecture engine and that the present invention is applicableto other gas turbine engines including direct drive turbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, withthe engine at its best fuel consumption—also known as bucket cruiseThrust Specific Fuel Consumption (“TSFC”). TSFC is the industry standardparameter of 1 bm of fuel being burned divided by 1 bf of thrust theengine produces at that minimum point. “Low fan pressure ratio” is thepressure ratio across the fan blade alone, without a Fan Exit Guide Vane(“FEGV”) system. The low fan pressure ratio as disclosed hereinaccording to one non-limiting embodiment is less than about 1.45. “Lowcorrected fan tip speed” is the actual fan tip speed in ft/sec dividedby an industry standard temperature correction of [(T_(ambient)°R)/518.7° R)^(0.5)]. The “Low corrected fan tip speed” as disclosedherein according to one non-limiting embodiment is less than about 1150ft/second.

Referring to FIG. 2, a core housing 60 includes an inlet case 62 and anintermediate case 64 that respectively provide an inlet case flowpath 63and a compressor case flowpath 65. In other embodiments, the corehousing may include additional cases. Similarly, the compressor sectionas a whole may include any number of cases. Together, the inlet andcompressor case flowpaths 63, 65, in part, define a core flowpaththrough the engine 20, which directs a core flow C.

The intermediate case 64 includes multiple components, including theintermediate case portion 66, and the bearing support 68 in the example,which are removably secured to one another. The bearing support portion68 has a first bearing 70 mounted thereto, which supports the innershaft 40 for rotation relative to the intermediate case 64. In oneexample, the first bearing 70 is a ball bearing that constrains theinner shaft 40 against axial and radial movement at a forward portion ofthe inner shaft 40. The first bearing 70 is arranged within a bearingcompartment 71.

In the example, the inner shaft 40 is constructed of multiple componentsthat include, for example, a main shaft 72, a hub 74 and a flex shaft76, which are clamped together by a nut 80 in the example. The firstbearing 70 is mounted on the hub 74 (i.e., low pressure compressor hub).The flex shaft 76 includes first and second opposing ends 82, 84. Thefirst end 82 is splined to the hub 74, and the second end 84 is splinedto and supports a sun gear 86 of the geared architecture 48. Bellows 78in the flex shaft 76 accommodate vibration in the geared architecture48.

The geared architecture includes star gears 88 arrangedcircumferentially about and intermeshing with the sun gear 86. A ringgear 90 is arranged circumferentially about and intermeshes with thestar gears 88. A fan shaft 92 is connected to the ring gear 90 and thefan 42 (FIG. 1). A torque frame 94 supports the star gears 88 andgrounds the star gears 88 to the housing 60. In operation, the innershaft 40 rotationally drives the fan shaft 92 with the rotating ringgear 90 through the grounded star gears 88.

The low pressure compressor 44 includes multiple compressor stagesarranged between the inlet and intermediate case flowpaths 63, 65, forexample, first and second compressor stages 98, 100, that are secured tothe hub 74 by a rotor 96. The first bearing 70 is axially aligned withone of the first and second compressor stages 98, 100. In one example, avariable stator vane array 102 is arranged upstream from the first andsecond compressor stages 98, 100. Struts 104 are arranged upstream fromthe variable stator vane array 102. An array of fixed stator vanes 106may be provided axially between the first and second compressor stages98, 100. Although a particular configuration of low pressure compressor44 is illustrated, it should be understood that other configurations maybe used and still fall within the scope of this disclosure.

The inlet case 62 includes inlet case portions 108, and bearing support110, which are removably secured to one another. The bearing supportportion 110 and torque frame 94 are secured to the inlet case portion108 at a joint 109. The bearing support portion 110 supports a secondbearing 112, which is a rolling bearing in one example. The secondbearing 112 is retained on the hub 74 by a nut 113, for example, and isarranged radially outward from the flex shaft 76 and radially betweenthe torque frame 94 and flex shaft 76. In the example, the secondbearing 112 is axially aligned with and radially inward of the variablestator vane array 102. The geared architecture 48 and the second bearing112 are arranged in a lubrication compartment 114, which is separatefrom the bearing compartment 71 in the example.

Referring now to FIG. 3 with continued reference to FIG. 1, the coreflow path of the example engine 20 begins at a core inlet 160 andextends through and past the low-pressure compressor 44. The core inlet160 has a radially inner boundary 162 and a radially outer boundary 166.

A core inlet stator 170 is located at or near the core inlet 160. Thecore inlet stator 170 attaches to a core case 174 at the radially innerboundary 162. The core inlet stator 170 attaches to an inlet case 178 atthe radially outer boundary 166. The core inlet stator 170 extendsradially across the core flow path C.

In this example, the radially inner boundary 162 is positioned a radialdistance D₁ from the axis A. The distance D₁, in this example, alsocorresponds to the radial distance between a root 164 of the core inletstator 170 and the axis A. In this example, the root 164 of the coreinlet stator 170 is radially aligned with the radially inner boundary162 of the core flow path C.

After flow moves through the core inlet 160, the flow moves through acompressor inlet 182 into the compressor section 24. In this example,the compressor section inlet 182 is an inlet to the low-pressurecompressor 44 of the compressor section 24. The compressor inlet 182extends from a radially inner boundary 186 to a radially outer boundary190.

Notably, a blade 198 of a rotor within the low-pressure compressor 44extends from a root 202 to a tip 206. The blade 198 is located at ornear the compressor inlet 182. The blade 198 part of a compressor rotorwithin a first stage of the compressor section 24. The blade 198 is thuspart of a first stage rotor, or a leading blade of the compressorsection 24 relative to a direction of flow along the core flow path C.

In some examples, the blade 198 represents the axial position where airenters the compressor section 24 of the core flow path C. The blade 198extends radially across the core flow path C.

The radially inner boundary 186 is positioned a radial distance D₂ fromthe axis A. The distance D₂, in this example, also corresponds to theradial distance between the root 202 of the blade 198 and the axis A. Inthis example, the root 202 is radially aligned with the radially innerboundary 186 of the core flow path C.

In the example engine 20, a preferred ratio range of the distance D₂ tothe distance D₁ spans from about 0.65 to about 0.9, which provides arelatively low profile core flow path contour. High profile flow pathcontours have greater differences between D₂ and D₁, and thus larger“humps” between the core inlet 160 and the compressor inlet 182. Highprofile flow path contours introduce discontinuities that undesirablydisrupt the airflow and undesirably add weight to the engine 20. Theratio range of about 0.65 to about 0.9 is made possible, in part, by theincorporation of the geared architecture 48 into the engine 20. The“hump” in this example is generally area 200. Additionally, since atleast one of the first and second bearings 70, 112 are packaged radiallyinward of the low pressure compressor 44, the distance D₂ may be largeras compared to bearing arrangements which have bearings axially offsetfrom the compressor section. Thus, the axially compact low pressurecompressor and bearing arrangement also minimizes discontinuities in theflow path contour while reducing the axial length of the engine.

Other characteristics of the engine having this ratio may include theengine 20 having a specific inlet flow of the low pressure compressor atcruising speeds to be between about 30 lb/sec/ft² and about 37lb/sec/ft². The specific inlet flow is the amount of flow moving intothe compressor section 24 and specifically, in this example, into acompressor inlet 182 and through the compressor section 24.

Another characteristic of the example engine 20 is that the cruisespeeds of the example engine are generally Mach numbers of about 0.7 toabout 0.9.

Yet another characteristic of the engine 20 is that a temperature at aninlet to the high-pressure turbine 54 may be from about 2,000° F.(1093.33° C.) to about 2,600° F. (1426.66° C.). Maintaining temperatureswithin this range balances good fuel consumption, low engine weight, andlow engine maintenance costs.

Yet another characteristic of the engine 20 is that a tip speed ofblades in a rotor of the low-pressure compressor 44 (a compressor rotor)may be from about 1,050 fps (320 m/s) to about 1,350 fps (411 m/s).

In this example, the geared architecture 48 of the engine 20 may have agear ratio of greater than about 2.3, or more specifically, a ratio offrom about 2.2 to about 4.0.

It should also be understood that although a particular componentarrangement is disclosed in the illustrated embodiment, otherarrangements will benefit herefrom. Although particular step sequencesare shown, described, and claimed, it should be understood that stepsmay be performed in any order, separated or combined unless otherwiseindicated and will still benefit from the present invention.

Although the different examples have specific components shown in theillustrations, embodiments of this invention are not limited to thoseparticular combinations. It is possible to use some of the components orfeatures from one of the examples in combination with features orcomponents from another one of the examples.

Although an example embodiment has been disclosed, a worker of ordinaryskill in this art would recognize that certain modifications would comewithin the scope of the claims. For that reason, the following claimsshould be studied to determine their true scope and content.

1. A gas turbine engine comprising: a core housing including an inletcase and an intermediate case that respectively provide an inlet caseflow path and an intermediate case flow path, a geared architecturearranged within the inlet case; a shaft providing a rotational axis; ahub operatively supported by the shaft; a rotor connected to the hub andsupporting a compressor section, the compressor section being arrangedaxially between the inlet case flow path and the intermediate case flowpath; a bearing mounted to the hub and supporting the shaft relative toone of the intermediate case and the inlet case; a core inlet includingthe inlet case flow path and having a radially inner boundary that isspaced a first radial distance from the rotational axis, and acompressor section inlet having a radially inner boundary that is spaceda second radial distance from the rotational axis, wherein a ratio ofthe second radial distance to the first radial distance is of about 0.65to about 0.9; and a fan section driven by the geared architecture thatis driven by the shaft that rotates a compressor rotor within thecompressor section.
 2. (canceled)
 3. The gas turbine engine according toclaim 1, wherein the radially inner boundary of the core inlet is at alocation of a core inlet stator.
 4. The gas turbine engine according toclaim 1, wherein the radially inner boundary of the compressor sectioninlet is at a location of a compressor rotor.
 5. A gas turbine enginecomprising: a core housing including an inlet case and an intermediatecase that respectively provide an inlet case flow path and anintermediate case flow path, a geared architecture arranged within theinlet case; a shaft providing a rotational axis; a hub operativelysupported by the shaft; a rotor connected to the hub and supporting acompressor section, the compressor section being arranged axiallybetween the inlet case flow path and the intermediate case flow path; abearing mounted to the hub and supporting the shaft relative to one ofthe intermediate case and the inlet case; a core inlet including theinlet case flow path and having a radially inner boundary that is spaceda first radial distance from the rotational axis, and a compressorsection inlet having a radially inner boundary that is spaced a secondradial distance from the rotational axis, wherein a ratio of the secondradial distance to the first radial distance is of about 0.65 to about0.9, wherein the radially inner boundary of the compressor section inletis at a location of a compressor rotor; and wherein the compressor rotoris a first stage rotor of a low-pressure compressor.
 6. The gas turbineengine according to claim 1, wherein the core inlet is an inlet to thecore housing.
 7. The gas turbine engine according to claim 1, wherein aninlet mass flow rate flux of the compressor section is configured to befrom about 30 lb/sec/ft² to about 37 lb/sec/ft² when the engine isoperating at a cruise speed.
 8. The gas turbine engine according toclaim 1, wherein a turbine inlet temperature of a high-pressure turbinewithin the engine is configured to be from about 2,000° F. to about2,600° F. when the engine is operating at a cruise speed.
 9. The gasturbine engine according to claim 1, wherein a tip speed of a bladearray in the compressor section during engine operation is configured tobe from about 1,050 fps to about 1,350 fps.
 10. (canceled)
 11. The gasturbine engine according to claim 11, wherein the geared architecturehas a gear reduction ratio of greater than about 2.3.
 12. The gasturbine engine according to claim 1, wherein the inlet case includes afirst inlet case portion defining the inlet case flow path, and abearing support portion removably secured to the inlet case portion, thebearing mounted to the bearing support portion.
 13. The gas turbineengine according to claim 1, wherein the intermediate case includes anintermediate case portion defining the intermediate case flow path, anda bearing support portion removably secured to the intermediate caseportion, the bearing mounted to the bearing support portion.
 14. The gasturbine engine according to claim 1, wherein the bearing is a ballbearing.
 15. The gas turbine engine according to claim 1, wherein thebearing is a first bearing and further comprising a second bearingsupporting the shaft relative to the other of the intermediate case andthe inlet case.
 16. The gas turbine engine according to claim 15,wherein the first and second bearings are arranged in separate sealedlubrication compartments.
 17. The gas turbine engine according to claim1, wherein the geared architecture is coupled to the shaft, and a fan iscoupled to and rotationally driven by the geared architecture.
 18. A gasturbine engine comprising: a core housing including an inlet case and anintermediate case that respectively provide an inlet case flow path andan intermediate case flow path, a geared architecture arranged withinthe inlet case; a shaft providing a rotational axis, wherein the gearedarchitecture is coupled to the shaft, and a fan is coupled to androtationally driven by the geared architecture; a hub operativelysupported by the shaft; a rotor connected to the hub and supporting acompressor section, the compressor section being arranged axiallybetween the inlet case flow path and the intermediate case flow path; abearing mounted to the hub and supporting the shaft relative to one ofthe intermediate case and the inlet case; a core inlet including theinlet case flow path and having a radially inner boundary that is spaceda first radial distance from the rotational axis, and a compressorsection inlet having a radially inner boundary that is spaced a secondradial distance from the rotational axis, wherein a ratio of the secondradial distance to the first radial distance is of about 0.65 to about0.9; and wherein the shaft includes a main shaft and a flex shaft, theflex shaft secured to the main shaft at a first end and including asecond end opposite the first end, wherein the geared architectureincludes a sun gear supported on the second end.
 19. The gas turbineengine according to claim 18, wherein the shaft includes a hub securedto the main shaft, and the compressor section includes a rotor mountedto the hub.
 20. The gas turbine engine according to claim 19, whereinthe geared architecture includes a torque frame supporting multiplecircumferentially arranged star gears intermeshing with the sun gear,the torque frame secured to the inlet case.
 21. The gas turbine engineaccording to claim 1, wherein the rotor supports multiple compressorstages, and the bearing is axially aligned with and radially inward ofone of the compressor stages.
 22. The gas turbine engine according toclaim 21, wherein the compressor section includes a variable stator vanearray.
 23. The gas turbine engine according to claim 17, comprising alubrication compartment, wherein the geared architecture is arranged inthe lubrication compartment.
 24. The gas turbine engine according toclaim 1, wherein the core housing includes a core inlet stator having astator root that is spaced a first radial distance from the rotationalaxis, and the compressor section includes a compressor blade having ablade root that is spaced a second radial distance from the rotationalaxis, wherein a ratio of the second radial distance to the first radialdistance is of about 0.65 to about 0.9.
 25. The gas turbine engineaccording to claim 24, wherein the stator root is radially aligned witha radially inner boundary of a core flow path through the gas turbineengine.
 26. The gas turbine engine according to claim 24, wherein theblade root is radially aligned with a radially inner boundary of a coreflow path through the gas turbine engine.
 27. The gas turbine engineaccording to claim 24, wherein the core inlet stator is positionedwithin an inlet to a core section of the gas turbine engine.